Turbine blades of the abovementioned generic type are heat-resistant components which are used in particular in turbine stages of gas turbine arrangements and, in the form of guide blades or rotor blades, are exposed to the hot gases directly leaving the combustion chamber.
The heat resistance of such turbine blades comes, on one hand, from the use of heat-resistant materials and, on the other hand, from a highly efficient cooling of the turbine blades which are directly exposed to the hot gases and which, in order that a coolant, preferably cooling air, can continuously flow through and act on them, have corresponding cavities that are connected to a coolant supply system of the gas turbine arrangement, which coolant supply system provides cooling air for the purpose of cooling all of the gas turbine components which are exposed to heat, i.e. in particular the turbine blades, when the gas turbine is in operation.
Conventional turbine blades have a blade root to which the blade airfoil is connected directly or indirectly in the radial direction, which airfoil has a concave-shaped pressure side wall and a convex-shaped suction side wall which are integrally connected in the region of the blade leading edge and between which is bounded an interspace that, for cooling purposes, is supplied with cooling air from the direction of the blade root. In this context, the term “in the radial direction” signifies the turbine blade extent in the mounted state in the gas turbine arrangement, which is oriented radially with respect to the axis of rotation of the rotor unit. In order to undertake the supply and distribution of cooling air in the interspace enclosed between the suction side wall and the pressure side wall, for an optimized cooling of the turbine blade, the interspace is provided with intermediate walls which run in the radial direction and which in each case separate cavities that are oriented radially inside the blade airfoil, some of which have fluidic connections. At suitable positions along the cavities, throughflow openings are provided in the suction side wall or the pressure side wall, in the region of the turbine blade leading edge and/or trailing edge or at the turbine blade tip, such that the cooling air can escape outward into the hot gas duct of the turbine stage.
A gas turbine blade which has been optimized with respect to cooling purposes is known from EP 1 319 803 A2, which provides for a multiplicity of radially oriented cooling duct cavities inside the turbine blade airfoil which are in each case fluidically connected in meandrous fashion and through which more or less cooling air flows depending on the thermal load on the various blade airfoil regions. It is particularly expedient to provide particularly efficient cooling for the region of the blade leading edge, which experiences the greatest flow exposure and heat exposure to the hot gases. To that end, a cavity extends internally in the longitudinal direction with respect to the blade leading edge, which cavity is delimited by the suction side and pressure side which unite at the blade leading edge and by an intermediate wall which connects the suction side and pressure side with one another internally, this cavity being supplied with cooling air from the direction of the blade root. The cooling air flowing through the cavity usually leaves in the region of the blade airfoil tip. Furthermore, in order to improve the transfer of heat between the blade airfoil wall and the cooling air flowing through the cavity, structures that swirl the cooling air flow are provided along the wall regions which enclose the cavity.
A further preferred cooling of the blade leading edge region of a turbine blade is described in U.S. Pat. No. 5,688,104. Along the blade leading edge there runs a cavity which, on one hand, is bounded by the suction side wall and pressure side wall which unite at the blade leading edge, and by an intermediate wall which rigidly connects the suction side wall and pressure side wall to one another inside the blade airfoil. The cavity running along the blade leading edge is supplied with cooling air which enters the cavity only through cooling duct openings provided in the intermediate wall. The straight intermediate wall is provided, in its radial longitudinal extent, with a multiplicity of individual throughflow ducts through which cooling air from an adjacent radial cooling duct enters the aforementioned cavity along the blade airfoil, in the direction of the blade leading edge, in the manner of an impingement cooling. In order to evacuate the cooling air introduced into the cavity, film cooling openings are provided along the blade leading edge, respectively oriented toward the suction side outer wall and pressure side outer wall, through which openings the cooling air introduced into the cavity is expelled, forming a film cooling respectively on the pressure side outer wall and suction side outer wall.
In order to improve the cooling effect, especially of the blade leading edge of a turbine blade, it is appropriate with the known cooling techniques to, on one hand, increase the supply of cooling air and, on the other hand, optimize the cooling mechanisms of the impingement cooling.
Turbine blades which, for the purpose of an optimized heat resistance in particular in the region of the blade leading edge, have the abovementioned cooling measures nonetheless often exhibit, in the blade leading edge region along the pressure side wall and suction side wall, fatigue symptoms which, in the final stages, become apparent by the formation of cracks. The reason for such crack formation lies in the apparition of thermomechanical stresses, within the suction side wall and pressure side wall in the blade leading edge region, which stem from large temperature differences between the blade leading edge exposed to the hot gases and the inner wall regions of the blade airfoil which are acted upon by the cooling air. In particular in the case of transient operating states of the gas turbine arrangement, such as those which arise in the turbine stage during startup or changes in load, temperature differences of approximately 1000° C. may occur between the blade leading edge exposed to the hot gases and the intermediate wall and inner wall sections which are acted upon by the cooling air. It is obvious that such great temperature differences give rise, within the suction side wall and pressure side wall along the blade leading edge, to considerable thermomechanical stresses which lead to considerable material loads, as mentioned above.